Turbine machines, such as turbofan gas turbine engines or land based turbine generators, typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and mixed with fuel in the combustor section for generating hot combustion gases. The hot combustion gases flow through the turbine section which extracts energy from the hot combustion gases to power the compressor section and in the case of turbine generators, drive the turbine power shaft.
Many turbine machines include axial-flow type compressor sections in which the flow of compressed air is parallel to an engine centerline axis. Axial-flow compressors may utilize multiple stages to obtain the pressure levels needed to achieve desired thermodynamic cycle goals. A typical compressor stage consists of a row of rotating airfoils (called rotor blades) and a row of stationary airfoils (called stator vanes).
One design feature of an axial-flow compressor section that affects compressor performance and stability is tip clearance flow. A small gap extends between the tip of each rotor blade airfoil and a surrounding shroud in each compressor stage. Tip clearance flow is defined as the flow of fluid between the rotor tip and an outer shroud from the high pressure side (pressure side) to the low pressure side (suction side) of the rotor blade. Tip clearance flow reduces the ability of the compressor section to sustain pressure rise, increases losses and may have a negative impact on stall margin (i.e., the point at which the compressor section can no longer sustain an increase in pressure such that the gas turbine engine stalls).
At the airfoil tip in the region where the airfoil and its boundary layer interact with the endwall boundary layer and the tip leakage flow, the aerodynamic loading tends to be higher than at the airfoil midspan. High aerodynamic loading results in higher turning deviation, larger losses and an increased likelihood of boundary layer separation. Bulk separation of the boundary layer on rotor tips is one mechanism for compressor stall.